This invention is related to articles such as gas turbine blades and vanes protected by thermal barrier coating systems, and, more particularly, to modification of the thermal barrier coating for improved erosion resistance.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of hot gas turns the turbine, which turns the shaft and provides power to the compressor. The hot exhaust gases then flow from the back of the engine, driving it and the aircraft forward.
The hotter the exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the exhaust gas temperature. However, the maximum temperature of the exhaust gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-base superalloys, and can operate at temperatures of up to 1900.degree.-2100.degree. F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes. The composition and processing of the materials themselves have been improved. Physical cooling techniques are used. In one widely used approach, internal cooling channels are provided within the components, and cool air is forced through the channels during engine operation. In another approach, a thermal barrier coating system is applied to the turbine blade or turbine vane, which includes internal cooling channels in its design, which acts as a substrate. The thermal barrier coating system includes a ceramic thermal barrier coating that insulates the component from the hot exhaust gas, permitting the exhaust gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component. Ceramic thermal barrier coatings usually do not adhere well directly to the superalloys used in the substrates, and therefore an additional layer called a bond coat is placed between the substrate and the thermal barrier coating. The bond coat improves the adhesion, and, depending upon its composition and processing, may also serve as a diffusion barrier to prevent oxidation and corrosion damage of the substrate.
The use of thermal barrier coating systems increases the operating temperature of the protected component, and its environmental resistance. Erosion, another damage mode found in the gas turbine engine, has consequently become more significant as the life-limiting cause of failure of the protected components. The thermal barrier coating systems have relatively poor resistance to erosion by small particles that are generated in, or pass through, the combustor. These small particles impinge against the protected component, and can either gradually wear through the thermal barrier coating (typically if the particles are small) or cause it to chip and flake away (typically if the particles are large). Once the thermal barrier coating is removed in localized regions, the component may fail because those regions may then be exposed to temperatures above the maximum acceptable service temperature of the component's material.
There is a need to improve the erosion resistance of thermal barrier coating systems. The approach to improving erosion resistance must be compatible with existing techniques and cannot add significantly to the weight of the component. The present invention fulfills this need, and further provides related advantages.